1. Field of the Invention
The present invention relates to wind tunnel and aircraft instrumentation and more particularly to instrumentation for determining flow stagnation points.
2. Description of Prior Art
Design of aerodynamic surfaces, including airfoils turbine blades, propellers, fan blades and other lift-producing surfaces requires the accurate determination of flow field characteristics, either through experimental or analytic methods. Design of particular airfoil characteristics, such as camber, platform, thickness and others, will control laminar to turbulent flow transition points, along the airfoil and therefore control aerodynamic performance. Likewise, once a design is fixed, flight conditions, such as angle-of-attack, lift and drag, can be determined by locating stagnation points and other flow phenomena. Of these phenomena, flow stagnation points are of prime importance.
Despite the importance of locating the stagnation points, accurate location is difficult using conventional techniques. Design of leading edge shapes, where it is necessary to determine flow stagnation points, has typically required extensive wind tunnel testing. Conventional means of determining flow stagnation points, using various flow visualization techniques or using multiple total pressure sensors, also require multiple tunnel runs, particularly if it is desired to accumulate data across several speed ranges and angle-of-attack ranges. In the case of rotating surfaces, such as surfaces of propellers or turbine blades, accurate determination of flow stagnation is an even more difficult and lengthy process.
Experimental techniques typically include the use of a series of pressure ports around a leading edge. Because of the very large pressure gradients near the stagnation point, accuracy requires a large number of very small holes and associated pressure tubing in the stagnation region. Physical constraints of a wing generally preclude this type instrumentation except in limited experimental set-ups. Even in experimental set-ups, the design of airfoil sensors is laborious.
Likewise, analytic methods have also been cumbersome. Understanding of laminar-to-turbulent flow transitions has generally been restricted to steady state, two-dimensional flows. A variety of modeling techniques have been used to further describe transition-related phenomena such as "unit Reynolds Number effect", "Receptivity", "Bypass Mechanisms", and different types of breakdowns of boundary-layer instability. Nonetheless, the models for the physical mechanisms involved in the transition process have been inadequate. The major problem is that most models "freeze" the flow field, in either time or space, resulting in incomplete or erroneous results.
Freezing the flow field is not appropriate, even for leading edge flow, since experimental observations confirm unsteadiness at the very inception of the boundary layer. The influence of the leading-edge geometry and freestream parameters on the instability of laminar boundary layers has been the subject of many theoretical and experimental investigations. Goldstein (1985) found that the Stokes shear wave and the streamwise velocity gradient had a significant effect on the generation of Tollmien-Schlichting waves at the leading edge. Hall (1985) considered the instability mechanisms for the flow around a torsionally oscillating cylinder and concluded that when the basic state has a steady component of the same order as the oscillating part, then there is a strong possibility of interaction between the instability mechanisms associated with the steady and unsteady components. Lighthill (1954) in his pioneering work on the response of laminar boundary layers to fluctuations in the stream velocity obtained expressions for critical oscillatory frequencies for Hiemenz and Blasius flows. The influence of these frequencies on the leading-edge laminar boundary layer subject to acoustic excitation was investigated by Leehey & Shapiro (1979) who concluded that the T-S mode could be excited significantly.
In all the above studies, fluctuations in the laminar boundary layer were assumed to be caused by either oscillations in incoming flow or by oscillations of the body. The invention herein is based on a new analytic model which postulates that airflow around a body is inherently oscillatory even in laminar regions, including around the leading edge or near the stagnation point, because of transmitted effects of downstream flow.